This invention relates to the gas turbine blades used in gas turbine engines and, more particularly, to selectively protecting portions of the gas turbine blades with a protective coating.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine. These components include the turbine vanes and turbine blades of the gas turbine, upon which the hot combustion gases directly impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1800-2100xc2x0 F. These components are subject to damage by oxidation and corrosive agents.
Many approaches have been used to increase the operating temperature limits and service lives of the turbine blades and vanes to their current levels, while achieving acceptable oxidation and corrosion resistance. The composition and processing of the base materials themselves have been improved. Cooling techniques are used, as for example by providing the component with internal cooling passages through which cooling air is flowed.
In another approach used to protect the hot-section components, a portion of the surfaces of the turbine blades is coated with a protective coating. One type of protective coating includes an aluminum-containing protective coating deposited upon the substrate material to be protected. The exposed surface of the aluminum-containing protective coating oxidizes to produce an aluminum oxide protective layer that protects the underlying surface.
Different portions of the gas turbine blade require different types and thicknesses of protective coatings, and some portions require that there be no coating thereon. The application of the different types and thicknesses of protective coatings in some regions, and the prevention of coating deposition in other regions, while using the most cost-efficient coating techniques, can pose difficult problems for gas turbine blades which have previously been in service and are undergoing repair. In many cases, it is difficult to achieve the desired combination of protective coatings and bare surfaces. There is a need for an improved approach to such coating processes to achieve the required selectivity in the presence and thickness of the protective coating in some regions, and to ensure its absence in other regions. The present invention fulfills this need, and further provides related advantages.
The present approach provides a technique for selectively protecting a gas turbine blade which has previously been in service, and is undergoing refurbishment and/or repair. In one application, the protective coating on the airfoil is rejuvenated, while the underside of the platform of the gas turbine blade is given a platinum aluminide coating. The present approach is cost effective, and is usable even with relatively small gas turbine blades.
A method for protecting a gas turbine blade which has previously been in service includes the step of providing the gas turbine blade which has previously been in service. The gas turbine blade has an airfoil, a dovetail, and a platform therebetween having a top surface and a bottom surface. In a usual case, the gas turbine blade has no protective coating on the bottom surface of the platform.
The gas turbine blade is first cleaned. The step of cleaning may include the steps of removing surface dirt, oxides, and corrosion products from the airfoil, and removing surface dirt, oxides, and corrosion products from the platform. Such cleaning may be accomplished by contacting the turbine blade to a weak acid bath, and thereafter grit blasting the turbine blade. In the cleaning, it is preferred that the existing coatings on the airfoil not be removed.
A precious-metal first layer is first deposited on at least an airfoil first-layer region of the airfoil to form an airfoil portion of the first layer, and at least a platform first-layer region of the platform to form a platform portion of the first layer. The precious metal of the first layer may comprise, for example, platinum, palladium, or rhodium, or alloys thereof, but is preferably platinum. The first deposition step is preferably accomplished by electrodeposition. The first deposition step usually includes first masking any surfaces that are not to have the precious-metal first layer deposited thereon. The precious-metal first layer is preferably first deposited to a thickness of from about 0.00008 to about 0.000125 inches.
A precious metal second layer is second deposited overlying at least part of the platform portion of the first layer to form a platform portion of the second layer, but not overlying the airfoil portion of the first layer. The precious metal of the second layer may comprise, for example, platinum, palladium, or rhodium, or alloys thereof, but is preferably platinum. The second deposition step is preferably accomplished by electrodeposition. The second deposition step usually includes the second masking of surfaces that are not to have the precious-metal second layer deposited thereon. The precious metal second layer is preferably deposited so that a total thickness of the precious-metal first layer and the precious-metal second layer is from about 0.00018 to about 0.00032 inches.
An aluminum-containing layer is third deposited, preferably by vapor phase deposition, overlying at least the airfoil portion of the first layer and the platform portion of the second layer. The gas turbine blade is heated to interdiffuse the aluminum and the precious metal, preferably at least in part concurrently with the third deposition step. An airfoil precious-metal aluminide coating thickness on the airfoil at a conclusion of the step of heating is about 0.001 inch greater than an airfoil precious-metal aluminide coating thickness at a conclusion of the step of cleaning. A platform precious-metal aluminide coating thickness on the platform at a conclusion of the step of heating is about 0.0025 inch greater than a platform precious-metal aluminide coating thickness at a conclusion of the step of cleaning (which is usually zero).
Stated alternatively, a method for protecting a gas turbine blade which has previously been in service comprises the steps of providing the gas turbine blade which has previously been in service, the gas turbine blade having an airfoil, a dovetail, and a platform therebetween having a top surface and a bottom surface, and cleaning the gas turbine blade. The method further includes depositing a precious-metal first layer on an airfoil first-layer region of the airfoil, depositing a precious metal second layer on at least part of the platform, wherein the precious-metal second layer is thicker than the precious-metal first layer, depositing an aluminum-containing layer overlying at least the precious-metal first layer and the precious-metal second layer, and heating the gas turbine blade to interdiffuse the aluminum and the precious metal.
The conventional practice has been not to coat the bottom surface or underside (i.e., the surface adjacent to the dovetail and remote from the airfoil) of the platform. The present approach not only refurbishes and rejuvenates the airfoil by adding a new platinum aluminide protective coating, but also provides a first-time platinum aluminide protective coating to the bottom surface of the platform (if there has not previously been a platinum aluminide protective coating on the bottom surface) or thickens an existing platinum aluminide protective coating on the bottom surface of the platform. The platinum aluminide protective coating added to the airfoil is thinner and with less platinum than the platinum aluminide protective coating on the bottom surface of the platform, due to the two-step platinum-deposition procedure. At the same time, the dovetail surfaces remain uncoated, a requirement for mating with the turbine disk.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.